This invention relates generally to gas turbine engines, and more specifically to turbine shroud assemblies used in gas turbine engines.
Gas turbine engines generally include, in serial flow arrangement, a high pressure compressor for compressing air flowing through the engine, a combustor in which fuel is mixed with the compressed air and ignited to form a high energy gas stream, and a high pressure turbine. The high pressure compressor, combustor and high pressure turbine are sometimes collectively referred to as the core engine. Such gas turbine engines also may include a low pressure compressor, or booster, for supplying compressed air to the high pressure compressor.
Generally, gas turbine engines operate more efficiently as combustion and exhaust temperatures increase. However, the operating temperature of the combustion gases is normally limited by the materials used to fabricate the hot-section components of the engine, such as the combustor and the turbine. To facilitate operating the engine at a higher operating temperature, at least some known turbine assemblies are coated with a thermal barrier coating (TBC). The TBC facilitates thermally, insulating the components from the combustible gases.
To facilitate maintaining turbine tip clearance, known shroud assemblies are masked prior to applying the TBC. However, masking of shroud segments is a time consuming process that is typically done by hand. Furthermore, during the TBC application process, the tape may blister and deform when exposed to the heat generated during the TBC application process.